This invention relates generally to the field of gas turbine engines, and more particularly to a ceramic matrix composite airfoil member for a gas turbine engine and a method of manufacturing the same.
Gas turbine engines are known to include a compressor section for supplying a flow of compressed combustion air, a combustor section for burning a fuel in the compressed combustion air, and a turbine section for extracting thermal energy from the combustion air and converting that energy into mechanical energy in the form of a shaft rotation. Many parts of the combustor section and turbine section are exposed directly to the hot combustion gasses, for example the combustor, the transition duct between the combustor and the turbine section, and the turbine stationary vanes, rotating blades and surrounding ring segments.
It is also known that increasing the firing temperature of the combustion gas may increase the power and efficiency of a combustion turbine. Modern high efficiency combustion turbines have firing temperatures in excess of 1,600 xc2x0 C., which is well in excess of the safe operating temperature of the structural materials used to fabricate the hot gas flow path components. Special super alloy materials have been developed for use in such high temperature environments, and these materials have been used with specific cooling arrangements, including film cooling, backside cooling and insulation.
Ceramic and ceramic matrix composite (CMC) materials offer the potential for higher operating temperatures than do metal alloy materials due to the inherent nature of ceramic materials. This capability may be translated into a reduced cooling requirement that, in turn, may result in higher power, greater efficiency, and/or reduced emissions from the machine. High temperature insulation for ceramic matrix composites has been described in U.S. Pat. No. 6,197,424 B1, which issued on Mar. 6, 2001, and is commonly assigned with the present invention. That patent describes an oxide-based insulation system for a ceramic matrix composite substrate that is dimensionally and chemically stable at a temperature of approximately 1600xc2x0 C. That patent also describes a stationary vane for a gas turbine engine formed from such an insulated CMC material.
Prior art ceramic turbine airfoil members may be formed with an associated shroud or platform member. The platform defines a flow path between adjacent airfoil members for directing the hot combustion gasses past the airfoil members. The platform is exposed to the same high temperature gas environment as the airfoil member and thus may be formed of a ceramic or CMC material. The platform and the airfoil members may be formed as separate components that are unconnected and are allowed to have relative movement there between. However, such designs may not adequately transfer aerodynamic torque loads from the airfoil to the platform attachments. Alternatively, the platform and the airfoil may be formed as separate components that are mechanically joined together, as illustrated in U.S. Pat. No. 5,226,789. Such mechanical joints must be robust and thus tend to be complicated and expensive.
Another alternative for joining the airfoil and the platform is to form the platform and the airfoil as a single integral part. Monolithic ceramic is readily moldable to form, but it is limited to small shapes and is insufficiently strain tolerant for robust designs. CMC materials incorporate ceramic fibers in a ceramic matrix for enhanced mechanical strength and ductility. However, conventional ceramic composite processing methods increase in complexity and cost in a complex three-dimensional component such as a turbine vane. U.S. Pat. No. 6,200,092 describes a turbine nozzle assembly having a vane forward segment formed of CMC material wherein the reinforcing fibers are specially oriented across the juncture of the airfoil and the platform members. Such special fiber placement in the airfoil-to-platform transition region presents a manufacturing challenge, especially with insulated CMC construction. Furthermore, for some CMC compositions, shrinkage during processing may result in residual stresses in complex shapes that are geometrically constrained. The airfoil-to-platform attachment area is one area where such stresses would arise. Additionally, load transfer between the airfoil and the platform results in interlaminar stresses in the fillet region where mechanical properties may be compromised.
The drying of the prepreg fabric restricts the lay-up time available in wet lay-up processes. For large complex shaped parts, such as an integrally formed airfoil/shroud vane, the required lay-up time may exceed the allowable exposure time for the prepreg. Consequently, some portions of the component may dry before others, resulting in possible shrinkage cracks and related problems. Furthermore, the consolidation of complex parts frequently requires pressure application in multiple directions, thus requiring complex tooling and consolidation challenges.
A method of manufacture for a vane component of a gas turbine is described herein as including: forming an airfoil member of a ceramic matrix composite material; forming a platform member of a ceramic matrix composite material; and forming an integral vane component by bonding respective joint surfaces of the airfoil member and the platform member. The method may further include: forming the airfoil member of a ceramic matrix composite material in a green body state; forming the platform member of a ceramic matrix composite material in a green body state; and urging the respective joint surfaces of the airfoil member and the platform member together at a firing temperature to form a sinter bond there between. The method may include densifying the sinter bond with a matrix infiltration process. The method may further include reinforcing the sinter bond with a fastener connected between the respective joint surfaces. Alternatively, the method may include bonding the respective joint surfaces of the airfoil member and the platform member with an adhesive.
A vane component for a gas turbine is described herein as including: an airfoil member formed of a ceramic matrix composite material; a platform member formed of a ceramic matrix composite material; and a bond between respective joint surfaces of the airfoil member and the platform member. The bond may be a sinter bond formed by urging the respective joint surfaces together in a green body state at a firing temperature. The component may further include a density-increasing material infused into the sinter bond by a matrix infiltration process. The component may include a fastener connected between the respective joint surfaces. The bond may be an adhesive bond. The component may include a mechanical fastener connected between the respective joint surfaces, or a ceramic matrix composite reinforcing member sinter bonded to the respective joint surfaces. The reinforcing member may be a generally U-shaped cross-section having opposed legs disposed on opposed sides of the respective joint surfaces.